The present invention relates generally to the field of rocket engines, and more particularly to an improved rocket engine combustion chamber design and method of making the same wherein the combustion chamber has a first smaller diameter film cooled surface portion adjacent to a propellant injector and steps suddenly outward to a second larger diameter portion at a position spaced away from the propellant injector, wherein the film cooling together with the sudden expansion of the diameter of the combustion chamber result in an exceptionally high degree of combustion efficiency.
The field of rocket science has advanced rapidly during the latter half of the twentieth century from its relatively primitive beginnings. Early rockets were essentially experimental, pilotless aircraft which were operated by crude control systems. The tremendous technological advances in rocket propulsion have been accompanied by similar advances in other essential fields such as electronics, inertial guidance and control systems, aerodynamics, and material sciences. As a result, rockets today are manufactured for a variety of purposes, ranging from military applications to carrying scientific instruments for use in gathering information at high altitudes, either within or above the earth""s atmosphere.
While such rockets may vary considerably both in application as well as size, they all include three essential components: a mission payload which is to be carried by the rocket, a guidance and control system, and a power source for propelling them. The first of these components is the guidance and control system, which controls the flight path of the rocket. The second of the aforementioned components is the mission payload, which, as mentioned above, may vary widely, varying from scientific instruments, to surveillance equipment, to explosive warheads.
It is the third of the three essential components of a rocket, namely the power source, which is the focus of the present invention. This power source is typically a self-contained rocket engine. Three different types of rocket engines have been predominantly utilized in the past, namely solid propellant systems, liquid bipropellant systems, and liquid or gaseous monopropellant systems. Solid propellant systems present several significant disadvantages not found in liquid bipropellant and monopropellant systems. For example, solid propellant systems are relatively heavy, have lower attainable exhaust velocities, and offer poor control of operating level in flight (throttleability).
Liquid bipropellant systems use an oxidizer and a fuel which are tanked separately and mixed in the combustion chamber. Typically, such liquid bipropellant systems use hydrazine or monomethylhydrazine as the fuel, and nitrogen tetroxide as the oxidizer. In some applications, bipropellant systems use gels instead of liquids. Liquid monopropellant systems typically also use hydrazine as a monopropellant fuel. Since liquid bipropellant systems are more widely used, the discussion which follows focuses on such systems.
The typical components of a liquid bipropellant propulsion system are the rocket engine, fuel tanks, and a vehicle structure to maintain these parts in place and connect them to the mission payload. The liquid bipropellant rocket engine itself consists of a main chamber for mixing and burning the fuel and the oxidizer, with the fore end occupied by fuel and oxidizer manifolds and injectors and the aft end comprising a nozzle. The oxidizer and the fuel are transferred from their respective tanks by pumps or may be pressurized by gas and are supplied to the injector manifold at a high pressure. The oxidizer and the fuel are then injected into the combustion chamber in a manner that assures atomization and mixing so that they may be efficiently reacted to produce thrust from the rocket engine.
Two problems which must be faced in the implementation of a rocket engine design are maximizing the efficiency of combustion and dealing with the problem of heat in the rocket engine. It will at once be appreciated by those skilled in the art that it is desirable to have a combustion efficiency approaching as close as possible to 100 percent. In addition, those skilled in the art will also realize that the hot gas chemical and temperature environment of a rocket engine, if left unchecked, may damage or destroy the combustion chamber during the desired lifetime of the rocket engine (which is typically in the range of 35,000-100,000 seconds).
The main combustion chamber of larger rocket engines typically use regenerative propellant cooling, in which the combustion chamber includes a coolant jacket through which liquid propellant (usually fuel) is circulated at rates high enough to allow the rocket engine to operate continuously without an excessive increase in the combustion chamber wall temperature. Smaller rocket engines instead use direct rejection of heat from the combustion chamber to the space environment by radiation heat transfer.
Effective cooling of a liquid rocket engine in the thrust range of 1 Newton to 10,000 Newtons is typically accomplished by using liquid or gaseous film cooling of the combustion chamber wall which establishes a stratified layer of low temperature fluid adjacent to the inner wall of the combustion chamber. This is accomplished by establishing a film cooling injection pattern and a main core injection pattern wherein the injectors provide a primary inner core of high temperature gases and a peripheral layer of low temperature unmixed and partially mixed propellant gases. The unmixed propellant used for the film cooling and partially mixed propellants must then be reacted in a rapid and efficient manner in order to provide a maximum specific impulse efficiency rocket engine.
Several patents which are relevant to the present invention may be reviewed as background information. These patents are U.S. Pat. No. 3,074,469, to Babbitt et al, U.S. Pat. No. 4,785,748, to Sujata et al., U.S. Pat. No. 4,915,038, also to Sujata et al., all of which are assigned to the assignee of the present invention, as well as U.S. Pat. No. 4,882,904, to Schoenman, and U.S. Pat. No. 4,936,091, also to Schoenman. U.S. Pat. No. 3,074,469, U.S. Pat. No. 4,785,748, U.S. Pat. No. 4,882,904, U.S. Pat. No. 4,915,038, and U.S. Pat. No. 4,936,091 are each hereby incorporated herein by reference.
It is accordingly one of the principal objectives of the present invention that it result in a rocket engine having a design and method of manufacture which provide a highly effective cooling mechanism which protects the combustion chamber from damage or destruction caused by high temperature conditions. It is a further objective of the present invention that it minimize or eliminate the reactions that take place between the incompletely reacted fuel and oxidizer products and the combustion chamber wall materials. It is a related objective of the present invention that it optimize the temperature gradients between the various components of the rocket engine to provide effective cooling and minimize structural and thermal stresses.
It is another of the principal advantages of the present invention that it enhance the combustion efficiency of the rocket engine to the maximum degree possible. It is accordingly an objective of the present invention that the rocket engine combustion chamber be of a design which promotes a complete mixing of the propellants such that they may be completely reacted within the combustion chamber. It is a related objective of the present invention that mixing of the main core of gas with the film cooling layer is accomplished after the need for the film cooling layer is no longer required, but before the unmixed and unreacted propellants leave the combustion chamber.
The stepped expansion combustion chamber rocket engine of the present invention must be of a construction which is both durable and long lasting, and it must also require no maintenance to be provided by the user throughout its operating lifetime. In order to enhance the market appeal of the stepped expansion combustion chamber rocket engine of the present invention, it should also be of relatively inexpensive construction to thereby afford it the broadest possible market. Finally, it is also an objective that all of the aforesaid advantages and objectives of the stepped expansion combustion chamber rocket engine of the present invention be achieved without incurring any substantial relative disadvantage.
The disadvantages and limitations of the background art discussed above are overcome by the present invention. With this invention, three key aspects are incorporated into the design of a rocket engine combustion chamber. The first two of these three key aspects are the use of film cooling and a stepped expansion combustion chamber, which together provide the heretofore mutually exclusive benefits of effective cooling of the combustion chamber and superior mixing of the fuel and oxidizer resulting in highly efficient combustion.
The combustion chamber thus consists of two portions, namely a first portion referred to herein as a precombustion chamber and a second portion referred to herein as a main combustion chamber. The precombustion chamber has a first diameter, and is located intermediate to the injector manifold assembly and the main combustion chamber, the latter of which has a second diameter larger than the first diameter. The precombustion chamber and the main combustion chamber are coaxial and adjacent to each other, such that the combustion chamber extends radially outward from the first diameter to the second diameter at the intersection between the precombustion chamber and the main combustion chamber in a step-wise manner. This construction is the derivation of the reference to a xe2x80x9csteppedxe2x80x9d combustion chamber.
The convergent throat and exhaust nozzle sections of the rocket engine form the remainder of the rocket engine. This section is formed in one single continuous assembly and is connected to the main combustion chamber at the end opposite the precombustion chamber.
The injector manifold assembly contains fuel and oxidizer manifolds which are located therein, as well as injectors communicating between the respective fuel and oxidizer manifolds and the interior of the stepped combustion chamber. The fuel manifold will be supplied with pressurized fuel, while the oxidizer manifold will be supplied with pressurized oxidizer. The injectors establish two spray patterns into the stepped combustion chamber, namely a main core injection pattern and a film cooling pattern. The main core injection pattern will provide a primary inner core of well mixed, high temperature combusting gases, while the film cooling pattern will provide an annular peripheral layer of low temperature unmixed and partially mixed propellant gases immediately adjacent to the interior surface of the precombustion chamber.
It will therefore be appreciated by those skilled in the art that only the injector manifold assembly and the precombustion chamber come in contact with the oxidizer and fuel and partially reacted combustion products at low temperatures. The precombustion chamber is effectively cooled by film cooling, and the sudden expansion process effectively mixes the remaining fuel and oxidizer, allowing them to combust completely. The injector manifold assembly and the precombustion chamber come in contact with the oxidizer and fuel, partially reacted combustion products, and, in some cases, decomposing fuel, but only at relatively low temperatures. The larger diameter main combustion chamber is subjected to higher temperatures, but only in the presence of the fully combusted propellants. Thus, the entire stepped expansion combustion chamber of the rocket engine of the present invention is protected from being simultaneously exposed to both corrosive partially reacted combustion products and high temperatures.
The use of the sudden expansion design in the stepped expansion combustion chamber of the rocket engine of the present invention also enhances the combustion efficiency by promoting mixing of the main core of gas with the film layer after the need for the cooling effect provided by the film layer is no longer required. The sudden expansion design is also effective in providing a flame-holding and recirculation zone to increase the chamber residence time of the unreacted and partially reacted gases. The resultant momentum of the main core gases is designed to impinge on the precombustion chamber wall just prior to the sudden change in diameter of the combustion chamber. The mixed core and film is thus effectively reacted in the recirculation zone which results from the sudden dimensional expansion.
The third key aspect of the stepped expansion combustion chamber rocket engine of the present invention is the use of a material for the inner surface of the precombustion chamber which has a high degree of thermal conductivity. The use of a material having a relatively high thermal conductivity for the precombustion chamber will serve to minimize the wall axial and circumferential temperature gradient between the portion of the combustion chamber at which the sudden expansion occurs and the face of the injector manifold assembly. The use of moderate amounts of propellant as a film coolant effectively cools the injector manifold assembly and eliminates the need for thermal isolation of the injector manifold assembly and the combustion chamber.
The use of a material for the inner surface of the precombustion chamber which has a high degree of thermal conductivity may be facilitated in two ways. First, the precombustion chamber itself may be made of a material having a high degree of thermal conductivity. Alternately, an inner precombustion chamber liner made of a material having a high degree of thermal conductivity may be used. The latter approach has an advantage in that the inner precombustion chamber liner can be fitted inside the precombustion chamber, with the precombustion chamber itself being made of a material which is better suited for assembly together with the injector manifold assembly and the main combustion chamber.
It may therefore be seen that the present invention creates a rocket engine having a design and method of manufacture which provide a highly effective cooling mechanism which protects the combustion chamber from damage or destruction caused by high temperature conditions. The present invention minimizes or eliminates the reactions that take place between the incompletely reacted fuel and oxidizer products and the combustion chamber wall materials. In a related aspect, the stepped expansion combustion chamber rocket engine of the present invention optimizes the temperature gradients between the various components of the rocket engine, thereby providing effective cooling and minimizing structural and thermal stresses.
The stepped expansion combustion chamber rocket engine of the present invention also enhances the combustion efficiency of the rocket engine to the maximum degree possible. The stepped expansion combustion chamber is of a design which promotes a complete mixing of the propellants such that they may be completely reacted within the combustion chamber. In the stepped expansion combustion chamber rocket engine of the present invention, mixing of the main core of gas with the film cooling layer is accomplished after the need for the film cooling layer is no longer required, but before the unmixed and unreacted propellants leave the combustion chamber.
The stepped expansion combustion chamber rocket engine of the present invention is of a construction which is both durable and long lasting, and which will require no maintenance to be provided by the user throughout its operating lifetime. The stepped expansion combustion chamber rocket engine of the present invention is also of relatively inexpensive construction to enhance its market appeal and to thereby afford it the broadest possible market. Finally, all of the aforesaid advantages and objectives of the stepped expansion combustion chamber rocket engine of the present invention are achieved without incurring any substantial relative disadvantage.